This disclosure relates generally to gas turbine engines and, more particularly, to cooling techniques for the airfoil sections of turbine blades and/or vanes of the engine. In particular, the present application is directed to an insert for use in convective cooling of the airfoils of the gas turbine engine which are exposed to high-temperature working fluid flow.
In general, gas turbine engines are built around a power core comprising a compressor, a combustor and a turbine, which are arranged in flow series with a forward (upstream) inlet and an aft (downstream) exhaust. The compressor compresses air from the inlet, which is mixed with fuel in the combustor and ignited to produce hot combustion gases. The hot combustion gases drive the turbine section, and are exhausted with the downstream flow.
The turbine drives the compressor via a shaft or a series of coaxially nested shaft spools, each driven at different pressures and speeds. The spools employ a number of stages comprised of alternating rotor blades and stator vanes. The vanes and blades typically have airfoil cross sections, in order to facilitate compression of the incoming air and extraction of rotational energy in the turbine.
High combustion temperatures also increase thermal and mechanical loads, particularly on turbine airfoils downstream of the combustor. This reduces service life and reliability, and increases operational costs associated with maintenance and repairs.
Accordingly, it is desirable to provide cooling to the airfoils of the engine.